CN117098908A

CN117098908A - Turbine ring assembly for a turbomachine

Info

Publication number: CN117098908A
Application number: CN202280024892.1A
Authority: CN
Inventors: C·贾罗塞; 奥勒列恩·盖拉德; 帕斯卡尔·塞德里克·塔巴林; 阿瑟·保罗·加布里埃尔·尼姆豪瑟尔; 克莱门特·埃米尔·安德烈·卡森
Original assignee: SNECMA SAS
Current assignee: Safran Aircraft Engines SAS
Priority date: 2021-03-30
Filing date: 2022-03-25
Publication date: 2023-11-21
Also published as: FR3121469B1; US20240068376A1; WO2022208007A1; FR3121469A1; EP4314493A1

Abstract

The application relates to a turbine ring assembly (1) comprising: ring segments (40) made of a ceramic matrix composite, each ring segment comprising first and second attachment tabs (42, 43) and a cavity (30) for circulating an air flow; -a metal support (5) comprising a first flange (52) and a second flange (53) bearing axially upstream against a second tab (43); a first metal shield (56) disposed upstream of the first flange, the first metal shield comprising an inner periphery (561) axially bearing against the first tab (42) downstream and an outer periphery (562) attached to the first flange (52); and air passage apertures (9 a,9 b) formed in the inner periphery of the first shroud (56) and/or the second flange, the apertures being configured to ensure passage of air flow from the cavity to the outside of the assembly (1).

Description

Turbine ring assembly for a turbomachine

Technical Field

The present application relates to the technical field of turbines, in particular for aircraft turbines. More specifically, the present application relates to a turbine ring assembly for a turbine that includes a plurality of ring sectors of ceramic matrix composite material and an annular metal support for the turbine ring.

Background

The prior art includes, inter alia, documents EP-A1-3865682, FR-A1-3056632, EP-A1-3173583, US-A1-2018/051591, EP-A1-3115559 and US-A1-2018/073391.

In general, turbines (particularly for aircraft) include, from upstream to downstream, fans, low pressure compressors, high pressure compressors, combustors, high pressure turbines, and low pressure turbines.

The high pressure turbine of the turbine comprises at least one stage comprising a turbine stator formed by an annular row of stationary straightening blades and an impeller rotatably mounted in a cylindrical or frustoconical assembly of ring sectors arranged end to end along the circumference and forming a turbine ring downstream of the turbine stator. In the case of all metal turbine ring assemblies, it is necessary to cool all elements of the ring assembly, in particular the turbine ring subjected to the most hot gas flow. Since the cooling air flow used is collected from the main air flow of the engine, this cooling has a significant impact on the engine performance. Furthermore, the use of metal for the turbine ring limits the possibility of increasing the temperature at the turbine, which will enable improved performance of the aeroengine.

To address these issues, it has been decided to fabricate turbine ring sectors from Ceramic Matrix Composite (CMC) materials and no metallic materials are required.

CMC materials have good mechanical properties, making the ceramic matrix composite suitable for use as structural elements, and the ceramic matrix composite maintains these properties at high temperatures. Thus, the use of CMC materials can reduce the cooling airflow required during operation, thereby improving the performance of the turbine. Furthermore, the use of CMC materials has the advantage of reducing the weight of the turbine and reducing the thermal expansion effects encountered by the metal components. Each sector of the turbine ring made of CMC material is fitted with attachment elements made of metallic material of the ring assembly and the annular support of the turbine ring, and these metallic attachment elements are also subjected to the hot gas flow. As a result, by reducing the operating cooling airflow of the turbine ring, the metal attachment elements in contact with the turbine ring are more exposed to the hot gas flow. In this case, it is the metal attachment element that is subjected to significant mechanical stresses.

Accordingly, there is a need to retrofit existing turbine ring assemblies with ring segments made of CMC materials, and in particular, by reducing the mechanical stresses experienced by metal components in contact with the CMC ring segments during turbine operation.

Disclosure of Invention

To this end, the application proposes a turbine ring assembly for a turbomachine of an aircraft, the ring assembly extending about an axis a and comprising:

ring sectors of ceramic matrix composite forming a turbine ring, each ring sector comprising a first attachment tab and a second attachment tab extending radially outwards from an upstream end and a downstream end of the ring sector, respectively, these first and second tabs defining between them a cavity for circulating a flow of cooling air F,

an annular metal support for a turbine ring, comprising a first and a second annular flange located upstream and downstream respectively, extending radially inwards and configured to retain a first and a second attachment tab of each ring sector, said second flange abutting axially against the second attachment tab upstream with respect to an orientation of a gas flow G intended to pass through the ring assembly along said axis a, and

-a first annular metal shroud arranged upstream of the turbine ring and the first flange, said first shroud comprising an inner periphery axially abutting against the first attachment tab towards the downstream and an outer periphery attached to the first flange.

According to the application, the ring assembly further comprises air passage apertures formed in the inner periphery of the first shroud and/or the second flange, the air passage apertures being configured to provide an air outlet from the cavity.

This configuration enables efficient cooling of the metal elements of the ring assembly that are exposed to the hot gas flow. The cooling system according to the application integrates apertures in the inner periphery of the first shroud and/or in the second flange. More specifically, the cooling air circulation cavity of each ring sector is supplied with an air flow (called ventilation and cooling air) coming from the compressor upstream of the ring assembly of the turbine. The air flow is preferably discharged from the cavity of each of the ring segments through the apertures of the first shroud and/or the second flange, thereby absorbing heat and thus cooling the metal elements of the ring assembly. This enables improved turbine performance because the CMC turbine ring and the first metal shroud and/or the second metal flange can be cooled at a minimum flow rate due to the air flow that collects upstream of the ring assembly.

It is therefore an advantage of the present application to provide a simple, highly reliable design for a ring assembly in a turbine that is low cost and of small overall size.

The turbine ring assembly according to the present application may include one or more of the following features, taken independently of each other or taken in combination with each other:

the apertures of the first shroud are oriented radially outwards from upstream to downstream and/or the apertures of the second flange are oriented radially inwards from upstream to downstream;

the aperture of the first shroud is also preferably oriented in the circumferential direction (with respect to axis a);

the aperture of the second flange is also preferably oriented in the circumferential direction (with respect to axis a);

the inner periphery of the first shield comprises a radial annular surface for abutment on the first attachment tab, and the aperture formed on the shield opens radially downstream to the outside of this surface;

-the second flange comprises an inner periphery having a radial annular surface for abutment on the second attachment tab, and the aperture formed on the second flange opens radially upstream to the outside of this surface;

-the air passage apertures are regularly spaced about said axis a;

-the air passage apertures are grouped together in a plurality of series of apertures, the circumferential spacing about said axis a between apertures in the same series of apertures being smaller than the circumferential spacing about said axis a between apertures of two consecutive series of apertures;

-each series of orifices comprises three to ten orifices;

the orifice is circular and/or elliptical;

-air passage apertures are formed in the first shroud and the second flange;

the second flange comprises a first portion, a second portion and a third portion interposed between the first portion and the second portion, the first portion and the third portion being separated by a shoulder, wherein an orifice formed on the second flange opens upstream and opens into the shoulder.

The application also relates to a turbine for a turbomachine of an aircraft, comprising at least one turbine stator formed by an annular row of fixed straightening blades, and an impeller mounted for rotation downstream of the turbine stator and inside a turbine ring of a ring assembly according to one of the characteristics of the application.

Each series of apertures formed on the first shroud may be located between two trailing edges of two consecutive fixed blades upstream of the turbine ring and/or each series of apertures formed on the second flange of the annular support may be located between two leading edges of two consecutive fixed blades downstream of the turbine ring.

The application also relates to a turbine, in particular for an aircraft, comprising at least one turbine ring segment assembly according to one of the characteristics of the application, or a turbine according to the application.

Drawings

Other features and advantages will appear from the following description of non-limiting embodiments of the application, with reference to the attached drawings, in which:

FIG. 1 is a partially schematic axial cross-sectional half view of a high pressure turbine of a turbine according to the prior art;

FIG. 2 is a schematic perspective view of a turbine ring assembly of a high pressure turbine according to the prior art;

FIG. 3 is a schematic axial cross-sectional view of a turbine ring assembly according to one of the embodiments of the present application;

FIG. 4 is a partial schematic perspective view of the ring assembly of FIG. 3, wherein the second downstream flange of the annular support includes air passage apertures according to a first configuration, according to a first embodiment;

FIG. 5a is a schematic perspective view of an upstream side of a first shroud of the ring assembly of FIG. 3 or 4, the first shroud including apertures according to a first configuration;

FIG. 5b is a schematic perspective view of the downstream side of the first shroud of FIG. 5 b;

FIG. 6a is a schematic perspective view of the downstream side of the second downstream flange of FIG. 4 including an orifice according to a first configuration;

FIG. 6b is a schematic perspective view of the upstream side of the second downstream flange of FIG. 6 a;

FIG. 7a is a partial schematic perspective view of an upstream side of a ring assembly wherein a first shroud includes air passage apertures according to a second configuration;

FIG. 7b is a partial schematic perspective view of the downstream side of the ring assembly of FIG. 7a, wherein the second downstream flange includes air passage apertures according to a second configuration;

FIG. 8a is a schematic perspective view of the upstream side of the second downstream flange of FIG. 7 b;

FIG. 8b is a schematic perspective view of the downstream side of the second downstream flange of FIGS. 7b and 8 a;

FIG. 9a is a schematic perspective view of the upstream side of the first shroud of FIG. 7 a;

fig. 9b is a schematic perspective view of the downstream side of the first shroud of fig. 7a and 9 a.

Detailed Description

In general, in the present application, the terms "longitudinal" and "axial" refer to the orientation of a structural element extending in the direction of a longitudinal axis. The longitudinal axis may coincide with the rotational axis of the engine of the turbine. The term "radial" refers to the orientation of a structural element extending in a direction perpendicular to the longitudinal axis. The terms "inner" and "outer" and "inner" and "outer" are used to refer to positioning relative to a longitudinal axis. Thus, a structural element extending along a longitudinal axis includes an inner surface facing the longitudinal axis and an outer surface opposite the inner surface of the structural element. In the present application, the terms "upstream" and "downstream" are conventionally defined with respect to the flow orientation of the gas flow in the turbine.

Turbines generally include, from upstream to downstream, fans, low pressure compressors, high pressure compressors, combustors, high pressure turbines, and low pressure turbines.

More specifically, fig. 1 shows a portion of a turbine 10 that extends along a longitudinal axis X and that includes, from upstream to downstream, a combustion chamber 1a, a High Pressure (HP) turbine 1b, and a low pressure (BP) turbine 1c. Each stage of one of the turbines 1b, 1c comprises an annular row of guide or stationary straightening blades 20, 20' and an impeller 3, alternately arranged in a known manner. An annular row of stationary blades 20 of the HP turbine 1b forms the turbine stator 2. The impeller 3 (or rotor) is rotatably mounted in a cylindrical or frustoconical assembly 1 according to the prior art configuration downstream of the turbine stator 2.

The assembly 1 comprises a plurality of ring sectors 40 arranged end to end in the circumferential direction and forming a turbine ring 4 enclosing the impeller 3. The turbine ring 4 is suspended from the turbine housing 6 by an annular support 5. The annular support 5 of the assembly 1 comprises, at its inner periphery, a first annular radial flange 52 and a second annular radial flange 53, respectively upstream and downstream, which are connected to each other by a cylindrical portion 51.

The annular support 5 also comprises a frustoconical (fig. 1) or annular (fig. 2) portion 54 extending upstream and outwards with respect to the axis X. This portion 54 is connected, on the one hand, at its radially inner end to the cylindrical portion 51 and, on the other hand, at its radially outer end to a radially outer annular flange 55 for attachment to a corresponding annular flange 65 of the turbine housing 6. The portion 54 of the annular support 5 defines an annular enclosure 50 of the chamber 1a with a frustoconical wall 58. Ventilation and cooling air is supplied to enclosure 50 through perforations 58a in frustoconical wall 58. Perforations 52a are formed in the first flange 52 of the annular support 5 to establish fluid communication between the enclosure 50 and the cooling air circulation cavity 30 of each ring sector 40. The cavity 30 is delimited externally by a wall 51 of the annular support. Arrow F represents the flow orientation of the cooling air flow, in particular from the compressor (not shown) of the turbine 10 supplying air to the combustion chamber 1a.

At the upstream and downstream ends of the ring sector 4, the ring sector comprises a first and a second attachment tab 42, 43 for attachment to a first and a second flange 52, 53, respectively, of the annular support 5.

The turbine assembly 1 is described in more detail with reference to fig. 2, which shows a half view of a radial cross section of a turbine assembly according to another configuration of the prior art. The turbine assembly 1 of fig. 2 may be assembled in the turbine 10 of fig. 1.

Thus, the ring assembly 1 extends about a longitudinal axis a that is substantially parallel to the axis X of the turbine 10. Arrow DA represents the axial direction of the turbine ring 4, and arrow DR represents the radial direction of the turbine ring 4. For simplicity of illustration, fig. 2 is a partial view of a turbine ring 4, which is in fact a complete ring. Arrow G indicates the flow orientation of the gas flow in the turbine 1b.

Each ring sector 40 has a cross-section substantially in the shape of an inverted greek letter "Pi" (Pi) in the plane defined by the axial direction DA and the radial direction DR. The cross section comprises an annular base 41, a first radial attachment tab 42 and a second radial attachment tab 43. The ring sectors may have a cross-section with a shape other than "pi", for example a "K" or "O" shape. The annular seat 41 comprises an inner surface 41a and an outer surface 41b opposite to each other in the direction DR of the ring 4. The inner surface 41a of the annular base 41 may be coated with a layer 44 of wear resistant material to define a flow conduit for the gas flow in the turbine.

First and second attachment tabs 42 and 43 extend radially outwardly from upstream and downstream ends 421a and 421b, respectively, of each ring sector. In the example shown in fig. 2, the first tab 42 and the second tab 43 protrude outwardly (in the direction DR) from the outer surface 41b of the annular base 41 and the upstream end 421a and the downstream end 421b of each ring sector 40. The first tab 42 and the second tab 43 extend over the entire width of the ring sector 40, i.e. over the entire arc of a circle described by the ring sector 40, or over the entire circumferential length of the ring sector 40.

As described above, the annular support 5 fixed to the turbine housing 6 includes:

a central cylindrical portion 51 extending in the direction DA and having an axis of rotation coinciding with the axis a of the turbine ring 4 when the central cylindrical portion and the turbine ring are attached together,

a first annular flange 52 and a second annular flange 53, respectively upstream and downstream, the first flange 52 and the second flange 53 extending radially inwards (with respect to the direction DR) from the inner surface 51a of the portion 51.

The first flange 52 includes a first free end 524 and a second opposite end 525, the second opposite end 525 being connected to the inner surface 51a of the portion 51.

The second flange 53 includes a first portion 531, a second portion 532, and a third portion 533 between the first portion 531 and the second portion 532. The first and third portions 531, 533 may form an inner periphery (relative to the direction DR) of the second flange 53, and the second portion 532 may form an outer periphery (relative to the direction DR) of the second flange 53. The first portion 534 includes a first free end 534 and the second portion 532 includes a second end 535 that is connected to the inner surface 51a of the portion 51. The first portion 531 extends between the first end 534 and the third portion 533, and the second portion 532 extends between the third portion 533 and the second end 535. The first and third portions 531, 533 are separated by a shoulder 537. In the example shown in fig. 2, the inner periphery of the first portion 531 (in particular the radial annular surface 536 of the first portion 531) is in contact with the second attachment tab 43 of the turbine ring 4. The thickness of the first and third portions 531, 533 is greater than the thickness of the second portion 532 to provide greater rigidity to the second flange 53 than the upstream portion, including particularly the first flange 52, to reduce axial leakage of the ring in the case of linear abutment.

Referring to fig. 2, the assembly 1 includes a second annular shroud 57 in addition to the first annular shroud 56. Two shields 56, 57 are removably attached to the first flange 52 of the annular support 5. The first shroud 56 and the second shroud 57 are arranged upstream of the turbine ring 1 with respect to the flow orientation G of the gas flow in the turbine. The first shroud 56 is arranged downstream of the second flange 57. The first shroud 56 is a unitary piece and the second shroud 57 may be divided into a plurality of annular sectors of the second shroud 57 or may be a unitary piece.

The first shield 56 has a first free end 564 and a second end 565 that is removably attached to the annular support 5, and more particularly to the first flange 52. Further, the first flange 52 has a first portion forming an inner periphery 561 (relative to the direction DR) and a second portion forming an outer periphery 562 (relative to the direction DR). The inner periphery 561 extends between the first end 564 and the outer periphery 562, and the outer periphery 562 extends between the inner periphery 561 and the second end 565. When the ring assembly 1 is installed, the inner periphery 561 of the first shroud 56 (particularly the radial annular surface 566 of the first shroud 56) bears against the first attachment tab 42 of each of the ring segments 40, and the outer periphery 562 bears against at least a portion of the first flange 52.

The second shield 57 has a first free end 574 and a second end 575 opposite the first end 574 and in contact with the cylindrical portion 51. The second end 575 of the second shroud 57 is also removably secured to the annular support 5, more specifically, to the first flange 52. The second shroud 57 further includes a first portion forming an inner periphery 571 and a second portion forming an outer periphery 572. An inner periphery 571 extends between the first end 574 and the outer periphery 572, and the outer periphery 572 extends between the inner periphery 571 and the second end 575.

The first and second shields 56, 57 are shaped to have inner peripheries 561, 571 spaced apart from each other and outer peripheries 562, 572 contacting each other, the two shields 56, 57 being removably attached to the first flange 52 by means of attachment screws 82 and nuts 83, the screws 82 passing through apertures 570, 560 and 520 provided in the outer peripheries 572 and 562 of the two shields 56, 57 and in the first flange 52, respectively.

To hold ring sectors 40, and thus turbine ring 4, in place with annular support 5, ring assembly 1 comprises, for each ring sector 40, two first axial pins 84 (with respect to direction DA) cooperating with first attachment tabs 42 and first shroud 56, and two second axial pins 86 (with respect to direction DA) cooperating with second attachment tabs 57 and second flange 53. For each corresponding ring sector 40, the inner periphery 561 of the first shroud 56 includes apertures for receiving two first pins 84, and the third portion 533 of the second flange 53 includes apertures configured to receive two second pins 86. For each ring sector 40, each of the first and second attachment tabs 42, 43 includes an aperture configured to receive the first and second pins 84, 56.

The annular support 5 also comprises radial pins 88 (with respect to the direction DR) which enable the ring 4 to be pressed in a determined manner in a lower radial position, i.e. towards the pipe. There is a gap between the axial pins 84, 86 and the holes in the ring to compensate for differential expansion between the metal and CMC component that occurs upon heating. The radial pins 88 cooperate in the direction DR with apertures formed in the cylindrical portion 51 of the annular support 5.

As previously described with reference to fig. 1, air from the compressor of the turbine is collected upstream of the combustion chamber 1a and enters (via perforations 58a, 52 a) the cooling air circulation cavity 30 of each ring sector 40. Thus, the cavity 30 supplies the air flow F to all ring sectors 40 and cools them.

Each ring sector 40 of the turbine ring 4 is made of a Ceramic Matrix Composite (CMC) material, while the first and second flanges 52, 53 and the first and second shrouds 56, 56 of the annular support 5 are made of a metallic material. This arrangement of the turbine ring assembly 1 in fig. 2 has many of the drawbacks mentioned in the background section above, in particular the risk of mechanical stress and embrittlement of the first metal shroud 56 and/or the second metal flange 53 exposed to the hot gas flow of the turbine.

The turbine ring assembly 1 of the present application may also be adapted for installation in a turbine 10 as shown in fig. 1. Fig. 3 to 9b show various embodiments of the assembly 1 according to the application.

The turbine assembly 1 according to the application comprises a ring sector 40 made of CMC material, a metallic annular support 5 and a first metallic shroud 56 and a second metallic shroud 57 as described above with reference to fig. 2. The turbine assembly 1 according to the application differs from the turbine assembly 1 according to the prior art (fig. 2) in the presence of air passage apertures 9a,9b formed in the first shroud 56 and/or the second flange 53 of the annular support 5. These apertures 9a,9b enable air flow F to pass from the cooling cavity 30 of each ring sector 40 to the outside of the turbine ring assembly 1 (fig. 3). Thus, this arrangement of the assembly 1 according to the application enables cooling of the first shroud 56 and/or the second flange 52 with a minimum flow rate of air flow (from the cavity 30) and/or enables preventing reintroduction of duct gas towards the first and second attachment tabs 42, 43.

More specifically, fig. 3 to 6b show a first embodiment of a turbine ring assembly 1 according to the application.

In fig. 3, the air passage aperture 9a is formed in the first shroud 56, and the air passage aperture 9b is formed in the second flange 53. In a first modification (not shown), the orifice 9a may be formed only in the first shield 56. In a second variant (not shown), the aperture 9b may be formed only in the second flange 53.

In particular, the aperture 9a is formed in the inner periphery 561 of the first shroud 56. The apertures 9a may be oriented in the circumferential direction of the ring assembly (relative to axis a). In the example shown in fig. 3, these orifices 9a are oriented radially outwards (with respect to the axis a or direction DA) from upstream to downstream. The aperture 9a is particularly downstream and opens radially outward for abutment against a radial annular surface 566 on the first attachment tab 42. This enables the air flow F to be directed from the cavity 30 towards the turbine stator 2 upstream of the ring sector 40.

The aperture 9b in the second flange 53 is preferably formed in the third portion 533 of the second flange 53. The apertures 9b may be oriented in the circumferential direction of the ring assembly (relative to axis a). In the example shown in fig. 3, these orifices 9b are oriented radially inwards (with respect to the axis a or direction DR) from upstream to downstream. The aperture 9b is particularly downstream and opens radially outward for abutment against the radial annular surface 536 on the second attachment tab 43. This also enables the air flow F to be directed from the cavity 30 towards the turbine stator 2 downstream of the ring sector 40. In this example, the orifice 9b opens upstream and opens into a shoulder 537 of the second flange 53. Alternatively (not shown), the orifice 9b may open upstream and open to a radial annular surface 538 of the third portion 533, which surface 538 does not abut on the second attachment tab 43 of the turbine ring 4.

According to this first embodiment, the apertures 9a formed in the first shroud 56 are evenly spaced about the axis a, as shown in fig. 5a and 5 b. The apertures 9b formed in the second flange 53 are also evenly spaced about the axis a as shown in figures 4, 6a and 6 b.

The apertures 9a,9b may be circular and/or elliptical.

The apertures 9a,9b may be three to ten for each ring sector 40. In the example shown in fig. 5a to 6b, each ring sector 40 has five apertures 9a, 9b.

Fig. 7a to 9b show a second embodiment of a turbine ring assembly 1 according to the application.

The turbine ring assembly 1 of the second embodiment differs from the turbine ring assembly 1 of the first embodiment in that the air passage apertures are arranged in the first shroud 56 and/or the second flange 53 of the annular support 5.

According to the second embodiment, the air passage apertures 9a,9b are combined together in a plurality of series of apertures per ring sector 40. Each series of apertures 9a may be formed on the first shroud 56 as shown in fig. 7a, 9a and 9b, and/or each series of apertures 9b may be formed on the second flange 53 as shown in fig. 7b, 8a and 8 b. In particular, the circumferential spacing about axis a between apertures 9a,9b in the same series of apertures 9a,9b is smaller than the circumferential spacing about axis a between two consecutive series of apertures 9a, 9b. The circumferential spacing refers to the distance between two consecutive orifices or two consecutive series of orifices having a similar profile, measured circumferentially with respect to the axis a.

In the example shown in fig. 7a and 7b, each series of orifices 9a is located on the first shield 56, in particular on a first predetermined area Za of the first shield 56.

Fig. 7b shows a series of apertures 9b located on the second flange 53, in particular on a second predetermined zone Zb of the second flange 53. The position of each series of orifices 9a,9b in the predetermined zone Za, zb may be constant or variable, depending on the dimensions of the turbine ring 4 and the annular support 5. Preferably, each series of orifices 9a,9b is located in a zone Za, zb of highest static pressure in the flow duct of the gas flow G.

Each series of orifices 9a,9b may comprise three to ten orifices. In the example shown, each series of orifices 9a,9b comprises five orifices 9b.

The application also relates to a turbine (in particular an HP turbine 1 b) comprising at least one turbine stator 2,2 'formed by an annular row of fixed blades 20, 20' and an impeller 3. The impeller 3 is rotatably mounted downstream of the turbine stator 2 and inside the turbine ring 4 of the ring assembly 1 according to the application.

When the turbine 10 includes a single annular row of stationary blades (FIG. 1), the blades 20 upstream of the ring sector 40 are formed on the HP turbine stator 2, and the blades 20 'downstream of the ring sector 40 are formed on the BP turbine stator 2'. Alternatively, when turbine 10 includes a plurality of HP turbine stators, blades 20 'downstream of ring sector 40 are formed on BP turbine stator 2' and/or on HP turbine stator 2. As an example (not shown), the first series of apertures 9a formed on the first shroud 56 are located between the trailing edges 22 of two consecutive blades 20 upstream of the ring sector 40 and correspond to the first zone Za. The second series of apertures 9b formed on the second flange 53 are located between the leading edges 21 of two consecutive blades 20' downstream of the ring sector 40 and correspond to the second zone Zb. The ring assembly 1 enables a specific tangential positioning of the orifices 9a,9b with respect to the blades 20, 20' upstream and downstream of the ring sector 40. This arrangement makes it possible to limit the reintroduction of the duct gas towards the ring assembly 1, in particular towards the first metal shield 56, the second metal flange 53 of the annular support 5 and also towards the first and second attachment tabs 42 and 43. This is because the gas flow G through between the trailing or leading edges of two successive blades is less disturbed (little or no turbulence) and therefore includes a higher static pressure than the gas flow G through between the leading and trailing edges of the same blade. Thus, arranging the apertures 9a,9b in the path of the undisturbed gas flow enables a rapid and unimpeded evacuation of the air flow F from the assembly 1, while cooling the first shroud 56 and/or the second flange 53.

The application also relates to a turbine 10, in particular for an aircraft, comprising at least one turbine ring assembly 1 according to the application. The turbine may be a turbojet or a turboprop.

Claims

1. A turbine ring assembly (1) for a turbine (10) of an aircraft, the ring assembly (1) extending about an axis (a) and comprising:

-ring sectors (40) of ceramic matrix composite forming a turbine ring (4), each ring sector (40) comprising a first and a second attachment tab (42, 43) extending radially outwards from an upstream and a downstream end (426 a, 426 b) of the ring sector (40), respectively, these first and second tabs (42, 43) defining between them a cavity (30) for the circulation of a cooling air flow (F),

-an annular metal support (5) for the turbine ring (4) comprising a first and a second annular flange (52, 53) respectively upstream and downstream, extending radially inwards and configured to retain the first and second attachment tabs (42, 43) of each ring sector (40), respectively, the second flange (53) abutting against the second attachment tab (43) axially upstream with respect to the orientation of the gas flow (G) intended to pass through the ring assembly (1) along the axis (a), and

a first annular metal shroud (56) arranged upstream of the turbine ring (4) and the first flange (52), the first shroud (56) comprising an inner periphery (561) axially abutting against the first attachment tab (42) towards the downstream and an outer periphery (562) attached to the first flange (52),

characterized in that the ring assembly (1) further comprises air passage apertures (9 a,9 b) formed in the inner periphery of the first shroud (56) and/or the second flange (53), these air passage apertures (9 a,9 b) being configured to provide an air outlet from the cavity (30).

2. The ring assembly according to claim 1, wherein the apertures (9 a) of the first shroud (56) are oriented radially outwards from upstream to downstream and/or the apertures (9 b) of the second flange (53) are oriented radially inwards from upstream to downstream; the apertures (9 a,9 b) are also preferably oriented in the circumferential direction.

3. The ring assembly according to claim 1 or 2, characterized in that the inner periphery (561) of the first shroud (56) comprises a radial annular surface (566) for abutment on the first attachment tab (42), and in that the aperture (9 a) formed on the shroud (56) opens radially downstream to the outside of this surface (566).

4. The ring assembly according to any one of the preceding claims, wherein the second flange (53) comprises an inner periphery having a radial annular surface (536) for abutment on the second attachment tab (43), and the aperture (9 b) formed on the second flange (53) opens radially upstream to the outside of this surface (536).

5. The ring assembly according to any one of claims 1 to 4, wherein the air passage apertures (9 a,9 b) are regularly spaced around the axis (a).

6. The ring assembly according to any one of claims 1 to 4, wherein the air passage apertures (9 a,9 b) are combined together by a plurality of series of apertures, the circumferential spacing about the axis (a) between apertures in the same series of apertures being smaller than the circumferential spacing about the axis (a) between apertures of two consecutive series.

7. The ring assembly according to the preceding claim, wherein each series of apertures comprises three to ten apertures (9 a,9 b).

8. The ring assembly according to any of the preceding claims, wherein the apertures (9 a,9 b) are circular and/or elliptical in shape.

9. The ring assembly according to any of the preceding claims, wherein the air passage apertures (9 a,9 b) are formed in the first shroud (56) and the second flange (53).

10. The ring assembly according to any one of the preceding claims, wherein the second flange (53) comprises a first portion (531), a second portion (532) and a third portion (533) between the first and second portions (531, 532), the first and third portions (531, 533) being separated by a shoulder (537), wherein an orifice (9 b) formed on the second flange (53) opens upstream and opens into the shoulder (537).

11. Turbine (1 b) for a turbine (10) of an aircraft, the turbine comprising at least one turbine stator (2, 2 ') formed by an annular row of fixed straightening blades (20, 20 ') and an impeller (3) mounted for rotation downstream of the turbine stator (2, 2 ') and inside the turbine ring (4) of the ring assembly (1) according to any one of claims 1 to 10.

12. Turbine according to claim 11 when dependent on any one of claims 6 to 10, characterized in that each series of orifices (9 a) formed on the first shroud (56) is located between two trailing edges (22) of two consecutive fixed blades (20) upstream of the turbine ring (4) and/or each series of orifices (9 b) formed on the second flange (53) of the annular support (5) is located between two leading edges (21 ') of two consecutive fixed blades (20') downstream of the turbine ring (4).

13. A turbine (10) comprising a turbine ring assembly (1) according to any one of claims 1 to 10 or a turbine (1 b) according to claim 11 or 12.